The present invention relates generally to gas turbine engines, and, more specifically, to a turbine nozzle therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through corresponding turbines. A high pressure turbine (HPT) immediately follows the combustor and extracts energy from the combustion gases for powering the compressor. A low pressure turbine (LPT) follows the HPT for extracting additional energy from the combustion gases for producing output work such as powering an upstream fan in an aircraft turbofan engine application.
The turbines include a stationary or stator nozzle having a row of nozzle vanes which direct the combustion gases into a corresponding row of turbine rotor blades extending radially outwardly from a supporting rotor disk. The nozzle vanes and corresponding rotor blades cooperate for extracting energy from the gases for in turn rotating the supporting rotor disk, which in turn is joined by a corresponding shaft to either the compressor rotor or fan rotor for rotating the corresponding blades thereof.
These engine components define an annular flowpath extending downstream between the corresponding rows of nozzle vanes and rotor blades. Since the nozzle vanes are stationary and the rotor blades rotate during operation suitable seals are required therebetween for confining the combustion gases in the intended flowpath for maximizing engine efficiency.
Since the flowpath confines the hot combustion gases, the various components defining that flowpath must be suitably cooled during operation to ensure a suitably long useful life. Since the first stage turbine nozzle of the HPT receives the hottest combustion gases it must be specifically configured for enhanced cooling operation, typically using a portion of compressor discharge air for the cooling thereof. And, since the remaining turbine nozzles are disposed downstream from the first stage nozzle the combustion gases are cooler, and the cooling requirements of those turbine nozzles are lower.
An exemplary turbofan aircraft engine enjoying many years of public use in this country includes a second stage HPT nozzle with a relatively simple cooling circuit therein. This reference turbine nozzle includes hollow nozzle vanes with imperforate pressure and suction sidewalls and a row of drilled holes along the trailing edge thereof. A three-pass serpentine channel is provided inside each vane, with an inlet in the radially outer band above the leading edge which receives compressor discharge air for internal cooling of the vane. The first channel of the serpentine circuit extends behind the leading edge to the inner band, and extends radially outwardly therefrom in a second or mid-chord channel which extends to the outer band, and then turns radially inwardly into a third and final channel which extends back to the inner band. A first outlet is provided through the inner band at the flow turn between the first and second channels, and a second outlet is also provided in the inner band at the end of the third channel. These two outlets discharge the spent cooling air from each vane into a purge cavity between the first stage rotor disk and the second stage nozzle to ensure effective cooling of the components in this region.
Since any air bled from the compressor is not used in the combustion process, such air decreases the overall efficiency of the engine and must be minimized. However, the bled air is nevertheless required to ensure suitable cooling of the various turbine components which require cooling for enhanced life.
Accordingly, the competing requirements for compressor air is a major design objective in gas turbine engines for enhanced life thereof. Cooling air must be bled from the compressor for ensuring long life. However, the cooling air must nevertheless be minimized to minimize the reduction in engine efficiency.
Another design objective is backflow margin. The cooling air leaves the compressor with maximum pressure and must be suitably channeled through the various turbine components which correspondingly reduces its pressure prior to being reintroduced back into the combustion flowpath. Pressure losses must be minimized to ensure a suitable backflow margin of the cooling air with a pressure suitably greater than the pressure of the combustion gases to prevent ingestion of those combustion gases into the cooled turbine components from which the cooling air is discharged.
The reference second stage turbine nozzle described above has enjoyed many years of successful commercial use. The nozzle is the subject of a development program for further enhancing its durability and longevity. As part of this development program the profile of the nozzle vanes themselves has changed in the leading edge region, which has correspondingly increased the need for cooling thereof for meeting higher life objectives. A significant problem, however, is that increasing cooling effectiveness in the leading edge region will correspondingly increase the temperature of the spent cooling air and decrease its effectiveness for cooling the downstream portions of the vane, as well as the turbine rotor components cooled by the purge air discharged from the second stage nozzle.
The leading edge flow channel in the reference nozzle is smooth without heat transfer enhancing turbulators therein and avoids the pressure drop in cooling air associated with the turbulators for maintaining effective backflow margin. Turbulators arc ubiquitous features in gas turbine engines found in various configurations and sizes in both turbine stator vanes and rotor blades. Although the primary objective of the turbulators is to increase heat transfer, and therefore increase cooling effectiveness of the limited cooling air bled from the compressor, turbulator performance varies from design to design, and between nozzle vanes or blades. Since turbine blades rotate during operation, the cooling air is subject to centrifugal forces which affects the heat transfer performance thereof, a phenomena not found in non-rotating nozzle vanes.
In a first stage turbine nozzle, performance of the turbulators must be maximized in view of the hottest combustion gases which bathe the nozzle vanes. Further cooling of the first stage nozzle vanes is typically also required, and is typically effected by introducing impingement baffles inside the nozzle vanes which first use the compressor discharge air in impingement cooling of the inner surfaces of the vane, prior to further cooling using turbulators therein.
However, second stage turbine nozzles are bathed in cooler combustion gases and may not require impingement baffles for their enhanced cooling capability, and may not require elaborate turbulator configurations for effective cooling. In fact, second stage nozzle vanes typically have imperforate pressure and suction sidewalls without the need for rows of film cooling holes often found in first stage nozzle vanes. The particular problem therefore in designing suitable second stage turbine nozzles is not maximizing cooling effectiveness with turbulators and impingement baffles, but balancing cooling of the various portions of the turbine nozzle while minimizing the amount of cooling air required therefor, and maintaining adequate backflow margin, and minimizing the rotor cooling air temperature.
Accordingly, it is desired to provide an improved second stage turbine nozzle for enhanced durability thereof.